Energy desaturation of electromechanical actuators used in satellite attitude control

ABSTRACT

In an attitude control system for a spacecraft having a frame which is to be maintained in a desired orientation relative to a selected set of reference axes, which system includes an electromechanical actuator mechanically coupled to the spacecraft frame for applying to the frame a mechanical torque which opposes deviations from the desired orientation relative to at least one reference coordinate, the actuator including an electric motor and an element connected to be rotated by the motor, and a torque command signal generator connected to the motor for applying a torque command signal which is a function of such deviations and which drives the motor in a manner to control the torque being produced by the actuator, there is included an energy desaturating system composed of: a unit coupled to the element for producing a signal representative of element motion; a signal modifying circuit connected to produce a control signal having a magnitude which is a function of the signal representative of element motion; and signal coupling components including a switch connected for applying the control signal to the actuator with a polarity for effecting actuator desaturation only during the times that the switch is closed.

BACKGROUND OF THE INVENTION

The present invention relates to attitude control systems employed inspace vehicles.

Space vehicles are generally provided with attitude control systems formaintaining the vehicle in a desired attitude with respect to some setof reference axes which do not rotate with the vehicle. In most cases,the reference axes either are aligned with the vehicle orbit or fixedwith respect to the sun or stars.

Conventional attitude control systems contain a plurality of channels,there generally being one channel associated with each vehicle bodyaxis, and each channel may contain an electromechanical actuator whichcan be in the form of a reaction wheel assembly, a momentum wheelassembly, or a control moment gyro assembly. Each type of assemblyincludes a wheel which is rotated at high speed to produce a forcetending to oppose disturbance forces acting on the vehicle.

During prolonged operation, such actuators tend to become saturated withunwanted disturbance energy.

In the case of reaction wheel assemblies, saturation occurs when thewheel drive motor reaches its limit speed.

Momentum wheel assemblies become saturated when the momentum wheel andthe space vehicle precess together under the influence of a disturbancetorque until the resulting angular displacement of the vehicle from adesired orientation becomes unacceptably large, or until the speed ofrotation of the momentum wheel deviates from the desired, fixed speed bya predetermined amount.

In the case of control moment gyros, saturation occurs when the wheelgimbal displacement becomes such that required control torque outputscan no longer be generated, or the gimbals reach the mechanical stopsthat limit their pivoting movement. Such gyro devices can have eitherone or two degrees of freedom. When the device has two degrees offreedom, the actuator includes two gimbal pivoting motors and two loops,or channels, each associated with a respective degree of freedom.

To desaturate a reaction wheel assembly, the usual practice is to applya voltage to the associated drive motor amplifier in a manner to reducethe speed of the reaction wheel. In the case of a momentum wheelassembly, a voltage is applied to the momentum wheel drive motoramplifier in order to return the momentum wheel speed to its nominalvalue. Moreover, torque is applied to the vehicle by means other thanthe momentum wheel assembly in order to return the vehicle and themomentum wheel to the desired attitude. For a control moment gyroassembly, desaturation is effected by acting on the or each gimbal motorto drive the associated gimbal toward its null position.

Whenever desaturation is being effected by applying voltage to a drivemotor amplifier, the result is to impose a disturbance torque on thespace vehicle. Therefore, it is usually necessary to provide anadditional source of stabilizing torque on the vehicle, at least duringthe desaturation phase.

Such an additional stabilizing source may be of the passive type, suchas gravity gradient booms or aerodynamic fins, or active devices not ofthe electromechanical variety, such as gas jet thrusters or magnetictorque rods.

It is the general practice in the art to effect momentum management, orenergy desaturation, by deriving signals representative of wheel motion.In the case of a reaction wheel, only a signal representative of wheelspeed is needed. If the actuator employs a momentum wheel, multiplesignals are needed: signals representative of the angular attitude ofthe satellite; and a signal representative of wheel speed. If theactuator employs a control moment gyro, the signal is representative ofthe angular position of a respective gimbal. The wheel motion signal isapplied to a computer in which a complicated algorithm is implemented toproduce a torque command voltage that will drive the wheel, or thegimbal, in the manner required to achieve desaturation.

All facets of space vehicle attitude control are described in a textedited by James R. Wertz, Spacecraft Attitude Determination and Control,D. Reidel Publishing Company, Dordrecht, Holland, 1978.

Various types of attitude control systems are described in the followingpublications:

V. N. Branets, et al., Development Experience of the Attitude ControlSystem Using Control Moment Gyros for Long-Term Orbiting Space Stations,38th Congress of The International Astronautical Federation, Brighton,United Kingdom, Oct. 10-17, 1987, pp 1-8;

H. F. Kennel, Steering Law for Parallel Mounted Double-Gimbaled ControlMoment Gyros--Revision A, NASA Technical Memorandum, TM-82390, January1981, available from NTIS.

Other publications describing attitude control system and energydesaturation, or momentum management, techniques are:

Bendix Research Laboratories, Southfield, Mich., Technical ReportBRL/TR-73-6768, CMG/TACS Control System Hybrid Simulation, September1973;

John R. Glaese, et al., Torque Equilibrium Attitude Control for SkylabReentry, NASA TM-78252, November 1979, available from NTIS;

Robert O. Hughes, Conceptual Design of Pointing Control Systems forSpace Station Gimballed Payloads, presented as Paper 86-1986 at the AIAAGuidance, Navigation, and Control Conference, Williamsburg, Va. Aug.18-20, 1986, pp 78-87;

Henry H. Woo, et al., Momentum Management Concepts for a Space Station,presented as Paper 86-2047 at the AIAA Guidance, Navigation, and ControlConference, Williamsburg, Va., Aug. 18-20, 1986, pp 277-286; a revisedversion was published in J. Guidance, Vol. 11, No. 1, January--February1988, AIAA, pp 19-25; and

Henry N. Woo, et al., Preliminary Evaluation of a Reaction ControlSystem for a Space Station, presented as Paper 86-2152 at the AIAAGuidance, Navigation, and Control Conference, Williamsburg, Va., Aug.18-20, 1986, pp 538-546.

The use of gravity gradient booms for spacecraft stabilization isdescribed in the following publications:

D. K. Anand, et al., Attitude Performance of Some Passively StabilizedSatellites, Journal of the British Interplanetary Society, Vol. 26,1973, pp 641-661;

David L. Blanchard, Flight Results from the Gravity-Gradient-ControlledRAD-1 Satellite, presented as Paper 86-2140 at the AIAA Guidance,Navigation, and Control Conference, Williamsburg, Va., Aug. 18-20, 1986,pp 479-487; and

R. V. Davis, et al., Flight Experience and Application of Earth-OrbitingGravity Gradient Stabilization Systems. (Proceedings of the SixteenthInternational Astronautical Congress, Anthens, Greece, InternationalAstronautical Federation, Vol. II, 1966), pp 293-300 and four pages ofFigures.

FIG. 1 illustrates one known type of attitude control system in whichspacecraft attitude is maintained by at least one auxiliary passivetorque source during actuator desaturation. In the drawing, mechanicalcouplings are illustrated by solid lines and electrical couplings bybroken lines, and one attitude control channel is illustrated.Spacecraft frame 2 is mechanically coupled to an electromechanicalactuator 4 and an auxiliary passive torque source 6. In normaloperation, actuator 4 will apply a torque tending to oppose deviationsin the attitude of frame 2 from its desired orientation relative to atleast one reference coordinate, or axis. This is indicated by thenegative sign in the line coupling actuator 4 to frame 2. Thus, innormal operation, actuator 4 tends to stabilize frame 2.

Auxiliary passive torque source 6 may, depending on the manner in whichthe satellite attitude is to be controlled, apply either a stabilizingor destabilizing torque to frame 2 during various orbital phases.

The true attitude angle and true angle rate of spacecraft frame 2 aresensed by angle motion sensors 8 which produce output signals indicativeof those parameters. Those signals are applied to a computer 10 whichderives, on the basis of those values, for a given attitude controlchannel, a torque command voltage which is applied to a power amplifier14 in actuator 4. The output voltage produced by amplifier 14 issupplied to an electric motor 16 to cause that motor to generate atorque which acts in a direction to return frame 2 to the desiredattitude about the coordinate axis associated with the attitude controlchannel.

The rotor of motor 16 is coupled to a momentum storage element 18 whichrotates together therewith. The motion of element 18 is indicative ofthe disturbance energy stored therein and is sensed by a sensor 20 whichsupplies a signal indicative of measured motion to computer 10.

If actuator 4 is a reaction wheel assembly or a momentum wheel assembly,element 18 is the wheel, which is driven by motor 16, and sensor 20produces a signal representative of the direction and speed of wheelrotation. If actuator 4 is a control moment gyro assembly, element 18 isa pivotally mounted gimbal supporting the gyro rotor and pivoted bymotor 16, and sensor 20 produces a signal representative of thedirection and magnitude of the angular deviation of the gimbal from itsnull position.

Computer 10 processes the information provided by the measured motionsignal according to a relatively complicated algorithm in order to applyto power amplifier 14 a signal which drives motor 16 in a manner todissipate the disturbance energy stored in element 18.

A system of the type illustrated in FIG. 1 would be employed, forexample, in an earth-oriented satellite, where the gradient of theearth's gravity is the available source of auxiliary passive restoringtorque with respect to the pitch and roll attitude control channels,while the inertial reaction of the spacecraft frame to the rotation ofthe satellite around the earth is the available source of auxiliarypassive restoring torque in the yaw channel.

In the case of a satellite which is orbiting the earth but which isoriented in a reference frame fixed in inertial space, for example asun-oriented or star-oriented satellite, the earth's gravity gradient isthe available source of auxiliary passive restoring torque for allattitude control channels. However, in this case, the ability of theauxiliary passive source of restoring torque to impose a stabilizinginfluence on the spacecraft frame is intermittent since, as thesatellite orbits the earth, the gravity gradient will alternatelygenerate restoring torque and disturbance torque. In this case,electromechanical actuator desaturation must be interrupted during thetime periods when the gravity gradient is generating a disturbancetorque because actuator desaturation at those times would destabilizethe satellite attitude.

FIG. 2 illustrates a second known type of attitude control system whichis identical to that shown in FIG. 1 except for the type of auxiliaryactuators employed. In the case of the arrangement shown in FIG. 2, theauxiliary actuators 22 are not passive torque sources, but instead areactuators of the active type, such as gas jet thrusters or magnetictorque rods. These actuators 22 would be controlled by appropriatesignals from computer 10. In the arrangement shown in FIG. 2,desaturation of the electromechanical actuators 4 would be effected inthe same manner as described above in connection with FIG. 1.

The desaturation techniques described above have a number ofdisadvantages which are connected with the fact that they must becontrolled by a digital computer operating according to a relativelycomplicated algorithm. Specifically, development, testing and correctionof the algorithm is a relatively time consuming, and hence expensive,procedure, particularly since they must allow for programmed attitudemaneuvers and must take account of the fact that the angle rates atwhich spacecraft rotate around their own centers of mass, in response totorques generated by the actuators, are finite. Moreover, when a newspacecraft is being developed, its performance parameters in space canonly be estimated. Once the spacecraft has been placed into orbit,adjustments to the algorithm employed to control desaturation, due todifferences between estimated and actual performance, are difficult, ifnot impossible, to effectuate.

Furthermore, the existing techniques raise the possibility of overshootin the torque commands to the electric motors driving the momentumstorage elements in the actuators which are being desaturated, resultingin unwanted energy being put back into the elements.

SUMMARY OF THE INVENTION

It is an object of the present invention to effect actuator desaturationin a manner which eliminates, or at least substantially reduces, theabove-described drawbacks of known techniques.

Another object of the invention is to control actuator desaturation by asimple feedback path which does not require a complicated algorithm forits operation.

A further object of the invention to effect actuator desaturation with acircuit arrangement which can be readily adjusted after the spacecrafthas been placed into orbit.

The above and other objects are achieved, according to the invention, inan attitude control system for a spacecraft having a frame which is tobe maintained in a desired orientation relative to a selected set ofreference axes, which system includes an electromechanical actuatormechanically coupled to the spacecraft frame for applying to the frame amechanical torque which opposes deviations from the desired orientationrelative to at least one reference coordinate, the actuator including anelectric motor and an element connected to be rotated by the motor, andtorque command signal generating means connected to the motor forapplying a torque command signal which is a function of such deviationsand which drives the motor in a manner to control the torque beingproduced by the actuator, by the provision of energy desaturating meanscomprising: means coupled to the element for producing a signalrepresentative of element motion; signal modifying means connected toproduce a control signal having a magnitude which is a function of thesignal representative of element motion; and signal coupling meansincluding a switch connected for applying the control signal to theactuator with a polarity for effecting actuator desaturation only duringthe times that the switch is closed.

The advantages of the present invention result essentially from the useof means connected to produce a motor control signal having a magnitudewhich is a function of the signal representative of wheel motion,together with a switch which selectively applies the motor controlsignal to the motor in order to effect actuator desaturation only duringthe times that the switch is closed.

The switch can be controlled according to principles already known inthe art either automatically by signals generated within the spacecraftor from the ground.

The signal modifying means can be a circuit device having a fixed gain,or can include, for example, a somewhat more complicated system, such asa cascade laglead compensator. However, at the present time, theadvantage of such more complicated system is not apparent. In any event,because of the simple nature of the signal modifying means, it could beconstructed in such manner that its gain can be varied under control ofsignals sent from the ground in order to adjust for differences betweenthe estimated, or design, and the actual performance of the actuator.

BRIEF DESCRIPTION OF THE DRAWINGS

FIG. 1 is a circuit diagram of one known type of attitude controlsystem, including one control channel.

FIG. 2 is a diagram similar to that of FIG. 1 of a second known type ofcontrol system.

FIGS. 1 and 2 have already been described above.

FIG. 3 is a diagram illustrating a system of the type shown in FIG. 1,modified in accordance with the present invention.

FIG. 4 is a diagram illustrating a system of the type shown in FIG. 2,modified in accordance with the present invention.

FIG. 5 is a diagram illustrating an analog embodiment of a fixed gaindevice which may be used in the attitude control systems of FIGS. 3 and4.

FIG. 6 is a diagram illustrating a digital embodiment of a fixed gaindevice which may be used in the attitude control systems of FIGS. 3 and4.

DESCRIPTION OF THE PREFERRED EMBODIMENTS

FIG. 3 illustrates a first type of system according to the presentinvention, which, as can be seen, differs from the arrangement as shownin FIG. 1 primarily with respect to the manner in which actuatordesaturation is controlled.

According to the invention, desaturation is effected by conducting ameasured element motion signal from sensor 20 to a device 24 having aselected gain, with the output of device 24 being conducted via a switch26 and a node which applies a negated version of the signal provided bydevice 24 to power amplifier 14 in order to drive motor 16 in adirection which will dissipate the unwanted energy which has been storedin the actuator.

During normal operation of actuator 4, when that actuator is operatingto maintain the attitude of frame 2 with respect to at least oneattitude axis, switch 26 is open so that the output of device 24 doesnot contribute to the control operation. When desaturation is to beeffected, switch 26 is closed, either by a signal generated undercontrol of programming within the vehicle or by a ground control signal.Then, the motion measured by sensor 20 is converted by device 24 into asignal which is applied with the requisite polarity to drive motor 16 ina manner which will effect desaturation.

As in the case of known systems, desaturation should be performed whenconditions are such that passive torque source 6 is producing arestoring torque since, during desaturation, actuator 4 will be applyinga destabilizing, or disturbance, torque to frame 2.

Because of the nature of the components 18, 20 and 24 in the feedbackpath, this feedback path has a low pass characteristic. This means thateven during desaturation, actuator 4 is capable of producing astabilizing response to short duration disturbance impulses.

The embodiment shown in FIG. 3 further differs from the correspondingprior art arrangement of FIG. 1 by the provision of a stabilizing torquecommand signal generating device 28 having a comparatively simple gaincharacteristic. Device 28 derives a stabilizing torque command signalfrom an appropriate combination of the measured attitude angle andmeasured angle rate signals associated with the respective referenceframe axis. Device 28 can be given relatively simple gain characteristicbecause the signal which it generates does not play a significant partin the desaturation operation. Depending on the characteristics of theparticular attitude control system involved, the gain of device 28 couldbe constant, frequency dependent, or time variable. Preferably, thegains of devices 24 and 28 are selected so that all of the poles of allof the attitude control channels are in the left half of the complexfrequency plane.

During desaturation, with switch 26 closed, the signal received by poweramplifier 14 is proportional to the difference between the torquecommand voltage generated at the output of device 28 and the negativefeedback voltage supplied by device 24, the gain of device 24 beingselected to assure that whenever switch 26 is closed, the voltagereceived by amplifier 14 will have a polarity sufficient to effect thedesaturation operation.

Embodiments of the type shown in FIG. 3 could be provided with auxiliarytorque sources which are constructed such that when the spacecraft is inits desired attitude and actuators 4 are not being desaturated, therestoring torque generated by the auxiliary actuators is in equilibriumwith disturbance torques acting on the spacecraft frame. In this manner,the control requirements imoposed on actuators 4 will be minimized. Theauxiliary torque sources could be constructed to establish this relationon the basis of principles well known in the art.

FIG. 4 illustrates an embodiment of the present invention in a systemhaving the form shown in FIG. 2. The embodiment shown in FIG. 4 differsfrom that of FIG. 2 not only by the provision of devices 24 and 28 andswitch 26, but additionally by the provision of a further signalgenerating device 30, preferably having a fixed gain, and a torquecommand switch 32 connected between device 30 and the actuator oractuators 22 associated with the particular attitude control channel.Since, in embodiments of this type, actuators 22 always produce arestoring torque, desaturation could be effected at any time. Theactuators 22 receive torque commands proportional in magnitude to themeasured deviation from the desired attitude angle based on the measuredattitude angle signal conducted from sensors 8. Device 30 will create asuitable sign inversion to assure that the torque command signal whichit produces has the polarity required to cause the associated actuators22 to produce restoring torques. Device 30 could have any one of thegain characteristics described above with reference to device 28.

Since actuators 22 are placed in operation only during desaturation, theswitches 26 and 32 of a given attitude control channel will be operatedin synchronism, so that both switches are open or closed at the sametime. In addition, these switches would be opened if the actuator oractuators 22 associated with a given channel should fail in a such amanner as to generate disturbance torques. If actuators 22 are gas jetthrusters, their operation only during desaturation procedures willserve to conserve the thruster fuel supply.

The desaturation technique characterizing the present invention isparticularly advantageous when applied to earth-oriented satellites,when the gravity gradient of the earth is used to generate the requiredauxiliary restoring torques in the roll and pitch channels of theattitude control systems, and the inertial reaction to the rotation ofthe satellite around the earth is used to generate the requiredauxiliary restoring torque in the yaw channel. In order to accomplishthis, constraints should be imposed on the mass distribution of thesatellite. These are the same constraints as would have to be imposed tomaintain attitude stability without active control.

In more general terms, desaturation can be achieved according to thepresent invention in any attitude control system other than one havingthe following combination of features: passive auxiliary restoringtorque sources; control moment gyro actuators; and orientation of thespacecraft in inertial space.

During a desaturation procedure, there is a decrease in the accuracywith which the spacecraft attitude is being controlled. Therefore, it ispreferable that desaturation be performed when some deviations inattitude can be tolerated.

If each actuator is associated with a given attitude axis, then theactuators associated with the various axes can be desaturated atdifferent times. On the other hand, if attitude control about each axisis effected by the combined operation of several actuators, then allactuators should be desaturated simultaneously. For example, there areattitude control systems which employ four actuators each of whichcontributes to the attitude control which respect to several axes.

The present invention permits desaturation to be controlled by arelatively simple device having a fixed gain, i.e. a gain which ispreferably independent of the signal supplied thereto, but which may, ifdesired, contain a lag-lead compensator. In any event, because of therelative simplicity of device 24, the invention permits the desaturationcircuitry for new satellites to be designed with far less effort thanrequired by the prior art approach. Moreover, device 24 can easily beimplemented as a separate, relatively simple circuit unit and this leadsto the possibility of constructing the unit so that its fixed gain canbe varied by signals transmitted from the ground after the satellite hasbeen placed into orbit. As a result, errors in initial designassumptions can be corrected after the satellite has been launched.

For this purpose, device 24 could be constituted by an analog circuitincluding an element such as a potentiometer whose slider can be shiftedby a suitable motor in order to vary the gain. Alternatively, device 24could be a digital device having an addressable memory composed of aplurality of columns each associated with a respective gain. Each columnwould contain a row of memory locations each having an addressassociated with a particular measured motion signal value and containinga representation of the output signal associated with that measuredmotion value. Then, in effect, the gain of the device would bedetermined by selection of an appropriate column by means of a signalsent from the ground. This digital technique is already well known inthe art, per se.

FIG. 5 illustrates, in simplified form, an analog embodiment of device24 in which the measured motion signal is applied to a differentialamplifier 40 having a preset gain via a potentiometer 42 which isconnected between the output of sensor 20 and the output of amplifier40. The slider of potentiometer 42 is electrically connected to thenegating input of amplifier 40 and mechanically coupled to a servo motor44 which is driven in response to signals transmitted from a groundstation in order to vary the gain of the circuit. The output ofamplifier 40 is connected to switch 26.

A digital embodiment of device 24 is shown in FIG. 6 and includes ananalog/digital converter and address decoder 50 which receives themeasured motion signal and converts that signal into a memory rowaddress representative of the amplitude of the measured motion signal. Asecond address decoder 52 generates a column address in response to asignal sent from a ground station. The resulting row and column addresssignals address a corresponding memory location in a memory matrix 54and the value stored at that location, which is representative of thedesaturation control signal to be produced, is supplied to adigital/analog converter 56 which then delivers a corresponding analogsignal to switch 26.

Device 30 of FIG. 4 could also have the form shown in one of FIGS. 5 and6.

While the description above shows particular embodiments of the presentinvention, it will be understood that many modifications may be madewithout departing from the spirit thereof. The appended claims areintended to cover such modifications as would fall within the true scopeand spirit of the present invention.

The presently disclosed embodiments are therefore to be considered inall respects as illustrative and not restrictive, the scope of theinvention being indicated by the appended claims, rather than theforegoing description, and all changes which come within the meaning andrange of equivalency of the claims are therefore intended to be embracedtherein.

What is claimed is:
 1. In an attitude control system for a spacecrafthaving a frame which is to be maintained in a desired orientationrelative to a selected set of reference axes, which system includes anelectromechanical actuator mechanically coupled to the spacecraft framefor applying to the frame a mechanical torque which opposes deviationsfrom the desired orientation relative to at least one referencecoordinate, the actuator including an electric motor and an elementconnected to be rotated by the motor, and torque command signalgenerating means connected to the motor for applying a torque commandsignal which is a function of such deviations and which drives the motorin a manner to control the torque being produced by the actuator, theimprovement comprising energy desaturating means comprising: meanscoupled to said element for producing a signal representative of elementmotion; signal modifying means connected to produce a control signalhaving a magnitude which is a function of the signal representative ofelement motion, said signal modifying means comprising a signalamplifying device having a fixed gain; and signal coupling meansincluding a switch connected for applying the control signal to saidactuator with a polarity for effecting actuator desaturation only duringthe times that said switch is closed.
 2. A system as defined in claim 1further comprising auxiliary torque means for applying to the frame anauxiliary torque which opposes deviations from the desired orientationwhen said switch is closed.
 3. A system as defined in claim 2 whereinsaid auxiliary torque means is constructed such that, when said energydesaturating means is inactive and the spacecraft is in its desiredattitude, the restoring torque generated by said auxiliary means is inequilibrium with disturbance torques acting on the spacecraft frame. 4.A system as defined in claim 1 wherein said actuator comprises areaction wheel assembly having a wheel which constitutes said element.5. A system as defined in claim 1 wherein said actuator comprises amomentum wheel assembly having a wheel which constitutes said element.6. A system as defined in claim 1 wherein said actuator comprises acontrol moment gyro having a rotor and a pivotally mounted gimbalsupporting said rotor, and said gimbal constitutes said element.
 7. Inan attitude control system for a spacecraft having a frame which is tobe maintained in a desired orientation relative to a selected set ofreference axes, which system includes an electromechanical actuatormechanically coupled to the spacecraft frame for applying to the frame amechanical torque which opposes deviations from the desired orientationrelative to at least one reference coordinate, the actuator including anelectric motor and an element connected to be rotated by the motor, andtorque command signal generating means connected to the motor forapplying a torque command signal which is a function of such deviationsand which drives the motor in a manner to control the torque beingproduced by the actuator, the improvement comprising energy desaturatingmeans comprising: means coupled to said element for producing a signalrepresentative of element motion; signal modifying means connected toproduce a control signal having a magnitude which is a function of thesignal representative of element motion; signal coupling means includinga switch connected for applying the control signal to said actuator witha polarity for effecting actuator desaturation only during the timesthat said switch is closed; and auxiliary torque means for applying tothe frame an auxiliary torque which opposes deviations from the desiredorientation when said switch is closed, said auxiliary torque meanscomprising at least one active component for generating the auxiliarytorque, auxiliary torque generating means for generating an auxiliarytorque command signal which is a function of such deviations, and secondsignal coupling means including a second switch connected for applyingthe auxiliary torque command signal to said active component only duringthe times that said second switch is closed.
 8. A system as defined inclaim 7 wherein said auxiliary signal generating means comprises anamplifying device having a fixed gain.
 9. A system as defined in claim 7wherein said switch and second switch are operated in unison.
 10. In anattitude control system for a spacecraft having a frame which is to bemaintained in a desired orientation relative to a selected set ofreference axes, which system includes an electromechanical actuatormechanically coupled to the spacecraft frame for applying to the frame amechanical torque which opposes deviations from the desired orientationrelative to at least one reference coordinate, the actuator including anelectric motor and an element connected to be rotated by the motor, andtorque command signal generating means connected to the motor forapplying a torque command signal which is a function of such deviationsand which drives the motor in a manner to control the torque beingproduced by the actuator, the improvement comprising energy desaturatingmeans comprising: means coupled to said element for producing a signalrepresentative of element motion; signal modifying means connected toproduce a control signal having a magnitude which is a function of thesignal representative of element motion, said signal modifying meanscomprising a circuit arrangement which produces an adjustable amplituderelation between the signal representative of element motion and thecontrol signal; and signal coupling means including a switch connectedfor applying the control signal to said actuator with a polarity foreffecting actuator desaturation only during the times that said switchis closed.
 11. A system as defined in claim 10 wherein said circuitarrangement comprises means responsive to a signal originating fromoutside the spacecraft for adjusting the amplitude relation.
 12. Asystem as defined in claim 11 wherein: said circuit arrangement isanalog circuit; said means responsive to a signal comprises a componenthaving a variable impedance; and said component is connected such thatthe adjustable amplitude relation is a function of the value of theimpedance and the impedance is varied by the signal originating fromoutside the spacecraft.
 13. A system as defined in claim 12 wherein saidcomponent comprises a potentiometer and said circuit arrangement furthercomprises a motor connected to adjust said potentiometer in response tothe signal originating from outside the spacecraft.
 14. A system asdefined in claim 11 wherein: said circuit arrangement is a digitalcircuit comprising: a memory having a plurality of groups of memorylocations, each group being associated with a respective amplituderelation and storing a succession of representations of control signalvalues corresponding to respective values of the signal representativeof wheel motion; and signal decoder means connected to said memory forderiving the control signal from representations stored in a respectivegroup of memory locations.
 15. A system as defined in claim 14 whereinsaid circuit arrangement further comprises means connected to saidmemory for selecting the group of memory locations from which thecontrol signal is derived in dependence on the value of the signaloriginating from outside the spacecraft.